1. Field of the Invention
This invention relates to a solid propellant grain suited for use in solid rocket motor assemblies and designed to reduce nozzle and insulation recession, especially for rocket motor assemblies comprising nozzle assemblies lined or made from carbon-based insulation material. This invention further relates to a solid rocket motor assembly containing the solid propellant grain.
2. Description of the Related Art
Conventionally, solid propellants of rocket motors contain, among other constituents, oxidizer and fuel components immobilized within a solid polymeric binder. The solid propellant is cast as a grain within a rigid outer case or shell of the rocket motor combustion chamber. A heat insulating layer and optionally a liner are usually interposed between the solid propellant grain and the outer case to protect the outer case from the high operating temperatures associated with rocket motor operation and to provide enhanced grain-to-case bonding.
The solid propellant grain is most commonly configured as either a center-perforated grain or an end burner grain. Generally, a center-perforated solid propellant grain includes a central perforation extending along a substantial portion of the length, if not the entire length, of the solid propellant grain and concentrically aligned with the longitudinal axis of the grain. As referred to herein, the cross section of the center perforation can have a relatively simple configuration, such as circular or oval; alternatively, the cross section can have a more complex geometry (star shaped, and/or with fins or slots) in order to increase exposed propellant surface area. The configuration of the center perforation geometry is usually tailored to provide a surface area corresponding to the desired rocket motor ballistic performance. Preferably, the centrally perforated propellant grain is ignited uniformly over its entire length, so that combustion reactions ensue simultaneously over the entire inner surface region. Oxidizing agents embedded within the solid-propellant grain drive these reactions to form large quantities of combustion products, which are expelled from the rocket motor through a nozzle in fluid communication with the combustion chamber. As the expelled combustion products thrust the rocket forward at high velocities, the combustion reactions continue to propagate outward radially to maintain the combustion of the propellant grain and expel combustion products through the nozzle.
An end burner solid-propellant grain functions in a similar manner to a center perforated grain, except that instead of a center perforation being present, the end burner solid-propellant grain is ignited at its aft end. The combustion reaction begins at the aft end and propagates along the length of the grain until reaching the forward end of the grain, at which point the grain stock is depleted. It is also known to use combinations of end burner and center perforated grains, in which the combustion reaction simultaneously occurs at both the aft end and center perforation of the propellant.
The amount of thrust produced by a rocket motor is proportional to the exit gas velocity squared. Nozzles are designed to accelerate the combustion product gases from the propellant grain to the maximum velocity at the exit. To achieve this end, nozzles usually have forward walls converging to a throat region and have aft walls diverging from the throat region to a larger exit area to form a converging-diverging nozzle configuration. The nature of compressible gases is such that a converging-diverging nozzle will increase the exit gas velocity and thereby thrust. It is within the purview of those skilled in the art to design a nozzle throat to optimize the ratio of exit area to throat area. Recession of the nozzle throat, however, decreases the thrust of an optimally designed nozzle.
As described above, an insulating layer and liner are typically placed between the propellant and the outer casing of the combustion chamber to protect the outer casing from the extremely high temperatures at which the rocket motor operates. Likewise, the nozzle must also be designed to withstand not only the elevated temperatures of the combustion products, but also the high velocities at which the combustion products pass over the nozzle inner surface.
Metal and metal alloys have been investigated for use as nozzle materials. One such heat resistant metallic material is silver-infiltrated tungsten. These materials provide generally satisfactory performance for large boosters. However, the weight penalty and expense associated with the presence of the tungsten and other suitable metals make such metal alloys impractical and uneconomical for many applications.
Carbon-based materials, such as carbon-carbon (C/C) composites and other graphitic materials, represent an excellent alternative to metals due to their high heat resistance, inexpensive cost, and relatively low weight.
It is widely acknowledged in the industry, however, that carbon-based nozzle throats tend to recede, especially at high operating temperatures and pressures. Several studies have identified oxidation reactions between the carbon-based nozzle material and oxygen-containing constituents of the combustion products as the dominant source of nozzle recession. As the propellant in the rocket motor burns, the carbon-based nozzle is exposed to the hot combustion gases, such as O.sub.2, H.sub.2 O, CO.sub.2, and NO. These gases, especially water and carbon dioxide, tend to react with the carbon-based nozzle materials to produce carbon monoxide gas, which is carried off with the discharged combustion products. As a consequence of the chemical consumption of the carbon-based nozzle materials, the nozzle throat recedes so that the nozzle throat passage area, which due to its narrow passage attributes to an increased velocity of combustion products, is undesirably increased.
The recession of the nozzle throat inner surface during motor operation is a source of several problems in rocket operation. One such problem is the performance loss encountered by the recession of carbon-based nozzle materials. As the nozzle throat material recedes, the accessible throat area for the combustion products to flow through increases, causing a corresponding decrease in output force. Additionally, rough nozzle surfaces, which tend to form during nozzle recession, have been shown to undergo recession at a faster rate than smooth surfaces. Thus, the nozzle throat recession process can be characterized as a self-perpetuating phenomenon. Another problem attributable to nozzle recession is a loss of predictability. Calculations for determining acceptable payloads and requisite propellant grain stocks must be accurate to ensure that the rocket will reach its intended target. The calculations necessary for ascertaining rocket dimensions and payloads are dependent upon many variables, including nozzle throat dimension. In-flight variations of nozzle throat dimension due to recession can significantly complicate, if not render impossible, precise motor performance calculations.
It would, therefore, be a significant improvement in the art to provide a rocket motor assembly that advantageously comprises an oxidizable nozzle material and/or insulation, such as carbon/carbon composite nozzle, yet in operation is capable of avoiding significant nozzle erosion, even when the operating conditions are characterized by high temperatures and pressures.